Present embodiments relate generally to gas turbine engines. More specifically, but not by way of limitation, present embodiments relate to composite airfoils having a metal patch structure located on one or both of an airfoil shank and dovetail.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is located at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, and a turbine. It will be readily apparent to those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure and high-pressure compressors, and low-pressure and high-pressure turbines. This, however, is not an exhaustive list.
The compressor and turbine generally include rows of airfoils that are stacked axially in stages. Each stage includes a row of circumferentially spaced stator vanes and a row of rotor blades which rotate about a center shaft or axis of the turbine engine. The turbine engine may include a number of stages of static air foils, commonly referred to as vanes, interspaced in the engine axial direction between rotating air foils commonly referred to as blades. A multi-stage low pressure turbine follows the two stage high pressure turbine and is typically joined by a second shaft to a fan disposed upstream from the compressor in a typical turbo fan aircraft engine configuration for powering an aircraft in flight.
An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades. The first and second rotor disks are joined to the compressor by a corresponding rotor shaft for powering the compressor during operation.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. The turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. The stator nozzles turn the hot combustion gas in a manner to maximize extraction at the adjacent downstream turbine blades. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy.
Due to extreme temperatures of the combustion gas flow path and operating parameters, the stator vanes and rotating blades in both the turbine and compressor may become highly stressed with extreme mechanical and thermal loading. Additionally, gas turbine engines often comprise turbofans which provide thrust. These turbofans also utilize airfoils to cause air movement from the forward toward the aft end of the engine and due to operating temperatures may be formed of lightweight composites.
One known means for increasing performance of a turbine engine is to increase the operating temperature of the engine, which allows for hotter combustion gas and increased extraction of energy. Additionally, foreign objects occasionally pass by these components with airflow. However a competing goal of gas turbine engines is to improve performance through weight reduction of components in the engine. One means of reducing weight of engine components is to reduce weight through the use of composite materials.
One desirable characteristic or design of gas turbine engines is to improve performance of airfoil structures. This may occur in a variety of fashions including use of composite materials. However, airfoils are often subjected to large centrifugal loads at steady state condition. The stress concentrations located at the upper end of the dovetail and lower end of the shank are often limiting factors for composite blade life and durability.
As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies with gas turbine engine components. More specifically, it would be desirable to overcome these deficiencies to improve life and durability of airfoils by reducing stress concentration and improving resistance to fatigue.
The information included in this Background section of the specification, including any references cited herein and any description or discussion thereof, is included for technical reference purposes only and is not to be regarded subject matter by which the scope of the invention is to be bound.